[21] In rocket
experiments, the measurement of the heat flowing from the combustion
gases to the engine walls and the use of this information to devise
satisfactory cooling of the engine are second only to obtaining
maximum performance. Without cooling, a flight-weight rocket engine
would be heated to its melting point in a second or two. Major
factors affecting this heat transfer are gas temperature, density,
and velocity; all three of these are much higher in rocket engines
than in other internal combustion engines. These factors, plus gas
composition, are functions of the propellants, engine design, and
operating conditions. The particular fuel and oxidizer, the
proportions used, combustion pressure, and combustion efficiency
determine gas composition, temperature, and density. Injector design,
propellant proportions, mass flow, and combustion chamber design
affect gas velocity. The rocket engineer seeks a design giving both
high performance and a cooling method for steady-state operation. He
is aided by combustion characteristics, for peak performance usually
occurs at a fuel-rich mixture where the heat transfer is lower than
at a leaner mixture.

[23] Heat transfer
measurements at Ohio State used two techniques common in rocket
experiments. In the "heat-sink" method, the combustion chamber and
nozzle are made from a high-conductivity material, usually copper, in
which a thermocouple to measure temperature is buried in the thick,
uncooled wall. During rocket operation, the high thermal conductivity
of the copper keeps the inside wall from melting as the heat rapidly
flows into the interior of the mass. This allows a rocket to operate
for a few seconds, and sometimes as long as 30 seconds. After the
run, the temperature of the copper mass comes to equilibrium and by
measuring this temperature, the total amount of heat absorbed can be
calculated from the known mass and specific heat of the copper. In
the second method, a water jacket surrounds comparatively thin engine
walls and a high-velocity water flow keeps the walls cool. The
average heat transfer can be obtained by measuring the water flow and
its temperature rise. Using these methods, Ohio State measured
average heat transfer rates of about 1.6 joules per second per square
meter (1 Btu/sec-sq in) for the combustion chamber and about twice
that for the nozzle. These values were on the same order as found in
high-performance rocket engines using other propellants, but are
several times higher than heat transfer rates in other types of
internal combustion engines and are, for example, from 20 to 200
times higher than in steam plants.

In mid-1948 a mechanical engineer from
Aerojet, Irwin J. Weisenberg, joined the Ohio State rocket staff
under Stary and specialized in heat transfer and cooling experiments.
The first attempt to use hydrogen as a coolant was to employ a porous
combustion chamber wall and force hydrogen through the wall into the
combustion chamber.10 This type of cooling, called transpiration or "sweat"
cooling, was popular at the time and work with it was under way at
several other rocket laboratories.

In the first part of 1949, another engineer at
the Ohio State rocket laboratory, Clair M. Beighley, made a
theoretical analysis in which a temperature ratio involving
combustion gas temperature, wall temperature, and coolant temperature
was related to dimensionless flow parameters. A porous combustion
chamber was tested later and the experimental data agreed with the
theoretical predictions. Porous wall chambers with uniform
permeability were difficult to make, however, and the Ohio State
rocket engineers turned to regenerative cooling when an analysis
showed it to be feasible. In this method, hydrogen is circulated in
coolant passages surrounding the engine prior to injection and
burning.

In the midst of preparations to try it
experimentally (in June 1949) Stary returned to Aerojet and still
another Aerojet engineer, Dr. Willard P. Berggren, arrived at Ohio
State as the new chief engineer for rocket
experiments.11

The experimental thrust chamber for
regenerative cooling was designed to produce 445 newtons at a chamber
pressure of 20.4 atmospheres (fig. 6).
Liquid hydrogen in the coolant jacket would be well above this value
and hence far above its critical pressure of 12.8 atmospheres so that
no boiling could occur in the coolant passages. The first successful
regenerative cooling run was on 26 August 1949, when the thrust
chamber operated for 60 seconds at an oxygen-to-hydrogen mass ratio
of 4.1 and produced an exhaust velocity of 3190 meters per
second-about 93 percent of theoretical performance.

In all, 33 successful runs were made, over
half of which operated for 60 or more seconds; one operated for 159
seconds. The runs covered a range of mixture ratios and...

[24] Fig. 6. Rocket
thrust chamber of 445 newtons designed to use liquid hydrogen-oxygen
and be regeneratively cooled by the liquid hydrogen, Ohio State
University, 1949. Scale and dimensions are inches. (Courtesy of 1. J.
Weisenberg.)

....the maximum exhaust velocity for the
series was 3270 meters per second.* In general, performance with the regeneratively-cooled
engine was considerably higher than that obtained with the
water-cooled chambers. The experimenters attributed this not only to
the elimination of heat losses, but also to a lower-density hydrogen
entering the combustion chamber, which produced improved mixing and
higher combustion efficiency. Figure 7
shows the regeneratively-cooled rocket operating in December 1949
during the series of tests. The frost on the chamber indicates that
it was well cooled.12

[25] Fig. 7. Liquid
hydrogen-oxygen rocket engine regeneratively cooled by the hydrogen.
Ohio State University, December 1949. Note the frost on the outside
of the rocket chamber and the shock diamonds in the exhaust.
(Courtesy of I.J. Weisenberg.)

* The highest performance run lasted 90 seconds at a
fuel-rich mixture (0/ F,4.7), 21 atm, and a relatively low overall
heat transfer rate of 2.1 J/s .
m2. In contrast, the
longest run (159 sec.) was at the stoichiometric mixture (0/ F,8),
19.6 atm, much lower exhaust velocity (2800 m/s), but almost triple
the overall heat transfer rate (5.2 J/s . m2). The comparison
illustrates that peak performance does not come at the same operating
conditions as maximum heat transfer. It also shows that hydrogen
cooling handled the higher heat load.